1. Field of the Invention
The present invention relates to turbomachinery and axial flow compressors. More particularly, the present invention pertains to a flow diverter or "scoop" which can be connected to the inner shroud region of a stator vane in an axial flow compressor of a gas turbine engine. The "scoop" comprises an annular foil which extends circumferentially around a rotor and is connected to the inner shroud region of a stator vane assembly in preselected stages of the compressor. The scoop intercepts leakage air flowing from the axially aft, high static pressure side of the stator vane assembly to the axially forward, low static pressure side of each stator vane assembly. The scoop re-directs this leakage air back into the working fluid flow such that a vector component of the re-directed air has an aftward velocity resulting in improved engine efficiency and stall margin.
2. Discussion of the Background
Gas turbine engines have been utilized to power a wide variety of vehicles and have found particular application in aircraft. The operation of a gas turbine engine can be summarized in a three step process in which air is compressed in a rotating compressor, heated in a combustion chamber, and expanded through a turbine. The power output of the turbine is utilized to drive the compressor and any mechanical load connected to the drive. Modern lightweight aircraft engines, in particular, have adopted the construction of an axial-flow compressor comprising a plurality of lightweight annular disk members carrying airfoils at the peripheries thereof. Some of the disk members are attached to an inner rotor and are therefore rotating (rotor) blade assemblies while other disk members depend from an outer casing and are therefore stationary (stator) blade or vane assemblies. The airfoils or blades act upon the fluid (air) entering the inlet of the compressor and raise its temperature and pressure preparatory to directing the air to a continuous flow combustion system. The stator vanes redirect and diffuse air exiting a rotating blade assembly into an optimal direction for a following rotating blade assembly. The air entering the inlet of the compressor is at a lower total pressure than the air at the discharge end of the compressor, the difference in total pressure being known as the compressor pressure ratio. Internally, a static pressure rise occurs across the stator vanes from diffusion and velocity reduction.
For a number of reasons having primarily to do with the design parameters of the cycle utilized in a particular engine, it is undesirable for the higher static pressure, higher static temperature air at the discharge side of a stator vane assembly to find its way back into the primary air flow at the inlet side of the stator vane assembly. This air, which returns to the relatively low static pressure area at the vane assembly inlet, is called leakage air and results in reduced engine efficiency. Particularly in the propulsion of aircraft, it is essential that the overall engine operate at a high efficiency level in order that the full advantages of the gas turbine engine may be realized. Leakage of air within the compressor thus detracts not only from the efficiency of the compressor itself but also the overall efficiency of engine operation.
Labyrinth seals connected radially inward from the stator vane assemblies of the compressor stage and connected to the inner rotor have long been utilized as a means to prevent leakage flow about the primary working fluid path around the stator vane assemblies. Notwithstanding the use of labyrinth seals, some leakage does occur, and this leakage air will travel, for example, from the high static pressure downstream side of a stator vane assembly to the lower static pressure at the upstream side of the stator vane assembly via a path which exists between the radial inward end of the stator vane assembly and the labyrinth seals connected to the rotor. After traveling to the upstream side of the stator vane assembly, the leakage air travels in a radially outward manner in the cavity existing between the stator vane assembly and adjacent rotor assembly. This radial path taken by the leakage air has a tendency to reduce the velocity and axial direction of air traversing the working fluid flow path of the compressor and tends to increase the amount of bleed air which further contributes to engine inefficiency.
Thus, a need is seen for a means for controlling leakage air flowing upstream and into the cavity existing between a stator vane and adjacent rotor blade and for preventing leakage air from impeding the forward momentum of air traversing the flow path of the compressor.